Casing assembly for a gas turbine engine

ABSTRACT

There is described a casing assembly for a gas turbine engine, the casing assembly comprising: an annular casing having a longitudinal axis; a unison ring assembly configured for rotation about the longitudinal axis; a plurality of vanes rotatably coupled to the annular casing; and a plurality of linkages, wherein each of the plurality of vanes is coupled to the unison ring assembly by a respective one of the plurality of linkages. Each linkage comprises: a first link coupled to the respective vane; and a second link rotatably coupled to the first link and rotatably coupled to the unison ring assembly, such that rotation of the unison ring assembly about the longitudinal axis effects rotation of the respective vane.

This disclosure claims the benefit of UK Patent Application No. GB1815599.4, filed on 25 Sep. 2018, which is hereby incorporated herein inits entirety.

The present disclosure relates to a casing assembly for a gas turbineengine

Existing gas turbine engines comprise compressor systems having aplurality of variable inlet guide vanes and a plurality of variablestator vanes. The variable inlet guide vanes and the variable statorvanes may be rotatable about their axes so as to optimise theperformance of the gas turbine engine. In order to achieve suchrotation, a casing assembly comprising an actuator, an actuator controlrod, a crankshaft, a plurality of stage control rods, a plurality ofbridge plates, a plurality of unison rings and a plurality of leversconnected in series to the variable inlet guide vanes and the variablestator vanes may be provided. However, such an arrangement is complex,expensive to manufacture and occupies a relatively large volume withinthe gas turbine engine.

It is therefore desirable to provide an improved casing assembly for agas turbine engine that overcomes these issues.

According to an aspect of the present disclosure there is provided acasing assembly for a gas turbine engine, the casing assemblycomprising: an annular casing having a longitudinal axis; a unison ringassembly configured for rotation about the longitudinal axis; aplurality of vanes rotatably coupled to the annular casing; and aplurality of linkages. Each of the plurality of vanes is coupled to theunison ring assembly by a respective one of the plurality of linkages.Each linkage comprises: a first link coupled to the respective vane; anda second link rotatably coupled to the first link and rotatably coupledto the unison ring assembly, such that rotation of the unison ringassembly about the longitudinal axis effects rotation of the respectivevane.

The unison ring assembly may be configured for movement on a singleplane.

The second link of each of the linkages may be rotatably coupled to itsrespective first link about a first axis of rotation that issubstantially perpendicular to the longitudinal axis.

The second link of each of the linkages may be rotatably coupled to theunison ring assembly about a second axis of rotation that issubstantially perpendicular to the longitudinal axis.

The first link of each of the linkages may be fixedly coupled to itsrespective vane.

The second link of each of the linkages may be rotatably coupled to itsrespective first link and/or the unison ring assembly by a hinge joint.

The second link of each of the linkages may be rotatably coupled to itsrespective first link and/or the unison ring assembly by a universaljoint.

The second link of each of the linkages may be rotatably coupled to itsrespective first link and/or the unison ring assembly by a sphericaljoint.

Two or more of the first links or two or more of the second links mayhave two or more different lengths.

The unison ring assembly may comprise two or more unison rings spacedfrom each other. The second links of each of the linkages may berotatably coupled to the two or more unison rings.

The two or more unison rings may be connected by one or more connectingrods.

Each of the two or more unison rings may comprise a plurality ofclevises. The second links of each of the linkages may be rotatablycoupled to the unison ring assembly via a respective one of theplurality of clevises.

Each of the two or more unison rings may comprise a plurality ofprojections extending towards the longitudinal axis. The annular casingmay comprise a plurality of grooves or slots for receiving the pluralityof projections and locating the two or more unison rings relative to theannular casing.

The two or more unison rings may be substantially planar.

The plurality of vanes may be a plurality of stator vanes and/or aplurality of inlet guide vanes.

A gas turbine engine may comprise a casing assembly as described in anypreceding statement.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm(around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390cm (around 155 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.3, 0.31,0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass duct may besubstantially annular. The bypass duct may be radially outside the coreengine. The radially outer surface of the bypass duct may be defined bya nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55 deg C.

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a perspective view of a casing assembly of a high pressurecompressor of the gas turbine engine;

FIG. 5 is a perspective view of a unison ring assembly of the casingassembly;

FIG. 6 is a close-up perspective view of a plurality of linkages of thecasing assembly;

FIG. 7 is a close-up perspective view of one of the linkages of thecasing assembly;

FIG. 8 is a cross-sectional view of the connection between a unison ringof the unison ring assembly and a casing;

FIG. 9 is a plan view of a plurality of linkages with the unison ringlocated at a first position;

FIG. 10 is a plan view of a plurality of linkages with the unison ringlocated at a second position;

FIG. 11 is a cross-sectional view of an alternative connection betweenan alternative unison ring and an alternative casing; and

FIG. 12 is a further cross-sectional view of the alternative connectionbetween the alternative unison ring and the alternative casing taken ina direction perpendicular to the cross-sectional view of FIG. 11.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38.

By way of further alternative example, the gearbox 30 may be adifferential gearbox in which the ring gear 38 and the planet carrier 34are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22meaning that the flow through the bypass duct 22 has its own nozzle thatis separate to and radially outside the core engine nozzle 20. However,this is not limiting, and any aspect of the present disclosure may alsoapply to engines in which the flow through the bypass duct 22 and theflow through the core 11 are mixed, or combined, before (or upstream of)a single nozzle, which may be referred to as a mixed flow nozzle. One orboth nozzles (whether mixed or split flow) may have a fixed or variablearea. Whilst the described example relates to a turbofan engine, thedisclosure may apply, for example, to any type of gas turbine engine,such as an open rotor (in which the fan stage is not surrounded by anacelle) or turboprop engine, for example. In some arrangements, the gasturbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 is a perspective view of a casing assembly 42 of a high pressurecompressor, such as the high pressure compressor 15 described withreference to FIGS. 1 and 2. The casing assembly 42 generally comprises acasing 44, a unison ring assembly 46, a plurality of linkages 48 and aplurality of vanes 50. Although only a single vane 50 is shown in FIG. 4for clarity, each of the linkages 48 is connected to a separate vane 50.The vanes 50 may be variable inlet guide vanes and/or variable statorvanes.

The casing 44 and the unison ring assembly 46 each have annularprofiles. The longitudinal axes of the annular casing 44 and the annularunison ring assembly 46 (i.e. the axes around which the casing 44 andthe unison ring assembly 46 extend) are aligned with a longitudinal axis41 of the casing assembly 42 (i.e. the axis around which the casingassembly 42 extends). The longitudinal axis 41 of the casing assembly 42is aligned with the principal rotational axis 9. Accordingly, the casing44, the unison ring assembly 46 and the casing assembly 42 extend aroundthe principal rotational axis 9. The unison ring assembly 46 isconfigured for rotation about the longitudinal axis 41. The unison ringassembly 46 is configured for movement on a single plane perpendicularto the longitudinal axis 41, and does not move in an axial directionalong the longitudinal axis 41.

The casing 44 is disposed within the unison ring assembly 46, such thatthe unison ring assembly 46 extends around the casing 44. The pluralityof vanes 50 are disposed within the casing 44, such that the casing 44extends around the vanes 50. The plurality of vanes 50 are disposedcircumferentially around the longitudinal axis 41. The plurality ofvanes 50 are arranged in three distinct rows (i.e. on three distinctplanes) disposed along the longitudinal axis 41. The plurality of vanes50 are each configured for rotation about respective axes 51 orientedperpendicularly to the longitudinal axis 41.

FIG. 5 is a perspective view of the unison ring assembly 46 inisolation. The unison ring assembly 46 comprises a first unison ring 52,a second unison ring 54 and a third unison ring 56. Each unison ring 52,54, 56 comprises an outer ring portion 58 and a series of alternatingclevises 60 and projections or centralisers 62 extending radiallyinwards from the outer ring portion 58 towards the longitudinal axis 41around the circumference of the unison rings 52, 54, 56. The projections62 and the clevises 60 are integrally formed with the outer ring portion58. Accordingly, each of the unison rings 52, 54, 56 may be cut out of asingle plate of metal using laser cutting or water jet cuttingprocesses.

The outer ring portions 58 are provided with a plurality of axiallyextending holes. The plurality of axially extending holes are disposedcircumferentially around the outer ring portions 58 and are spaced attwo distinct radial distances from the longitudinal axis 41.Accordingly, the axially extending holes are arranged in pairs ofconcentric circles. A plurality of identical threaded connecting rods 64extend through corresponding axially extending holes of the first,second and third unison rings 52, 54, 56 and are fixed to each of thefirst, second and third unison rings 52, 54, 56 by bolts. Accordingly,the first, second and third unison rings 52, 54, 56 are secured togetherto form a squirrel-cage structure that provides good out-of-planestiffness. The first, second and third unison rings 52, 54, 56 aresubstantially identical, with the exception of a tab 66 that extendsoutwards from the outer ring portion 58 of the second unison ring 54 andwhich contains a hole for direct connection to an actuator.

FIG. 6 is a close-up perspective view of the casing assembly 42 showingone of the vanes 50 and its respective linkage 48 in detail. The casing44 is provided with a plurality of through holes 68 directed towards thelongitudinal axis 41. The vane 50 extends through the through hole 68such that a radially outer end thereof is disposed outside of the casing44 and such that an aerofoil portion 70 thereof is disposed within thecasing 44. The longitudinal axis of the vane 50 is aligned with the axis51. The vane 50 is able to rotate within the through hole 68 about theaxis 51. The linkage 48 comprises a first link or lever 72 and a secondlink 74. A first end of the first link 72 is fixedly coupled to theradially outer end of the vane 50. A second end of the first link 72 ispivotally coupled to a first end of the second link 74. A second end ofthe second link 74 is pivotally coupled to the clevis 60. The first andsecond links 72, 74 are manufactured using a metal injection moulding(MIM) process.

FIG. 7 is a further close-up perspective view showing one of thelinkages 48. As shown, the radially outer end of the vane 50 is providedwith a shoulder 75 against which the first end of the first link 72 issecured by a bolt 76. Accordingly, the first link 72 and the vane 50 areunable to move relative to each other. The second end of the first link72 comprises a clevis 78 formed by two arms separated by a gap. Each ofthe arms comprises a through hole, through which a first pin 80 extends.The first pin 80 is rotatable within the through holes about an axis 82that is substantially oriented towards and perpendicular to thelongitudinal axis 41. The first end of the second link 74 comprises anotch within which the first pin 80 is located. A second pin 84 extendsthrough the second link 74 and the first pin 80 so as to secure thesecond link 74 to the first pin 80. The second link 74 is therefore ableto pivot relative to the first link 72 about the axis 82. The connectionbetween the second link 74 and the first link 72 is a hinge joint thatallows motion on a single plane. However, as outlined below, it mayalternatively be a different type of joint such as a universal joint ora spherical joint formed by a spherical bearing.

Wth reference to both FIGS. 6 and 7, the second end of the second link74 comprises a clevis 86 formed by two arms separated by a gap. Each ofthe arms comprises a through hole, through which a third pin 90 extends.The third pin 90 is rotatable within the through holes about an axis 92that is substantially oriented towards and perpendicular to thelongitudinal axis 41. The clevis 60 of the unison ring 52 is likewiseformed by two arms separated by a gap. Each of the arms comprises athrough hole through which a fourth pin 94 extends. The fourth pin 94 isrotatable within the through holes about an axis 96 extendingcircumferentially with respect to the unison ring 52. The fourth pin 94is fixedly connected to the third pin 90. The second link 74 istherefore rotatable about the axes 92 and 96. The connection between thesecond link 74 and the unison ring 52 is a universal joint allowingrelative movement of the two components about multiple axes.

FIG. 8 is a cross-sectional view of the connection between the unisonring 52 and the casing 44. The casing 44 comprises a first flange 98 anda second flange 100. The first and second flanges 98, 100 are spacedapart from each other so as to define a groove or slot 102. Alow-friction pad or coating 103 is disposed at the base of the groove102. The first and second flanges 98, 100, the groove 102 and thelow-friction pad 103 extend around the circumference of the casing 44(see FIGS. 4 and 6). The projection 62 is located within the groove 102,is supported by the low-friction pad 103 and is able to slide along thegroove 102 in a circumferential direction. Accordingly, the unison ringassembly 46 can be rotated about the longitudinal axis 41 (i.e. in acircumferential direction) via actuation of the actuator. The first andsecond flanges 98, 100 prevent the projection 62 and thus the unisonring 52 and the unison ring assembly 46 as a whole moving in an axialdirection, along the longitudinal axis 41 and along the casing 44.Further, since a plurality of projections 62 are provided that aredisposed around the circumference of the unison rings 52, 54, 56 (i.e.around the principal longitudinal axis 41), the unison ring assembly 46is centred with respect to the casing 44. Accordingly, deformation ofthe unison rings 52, 54, 56 (e.g. from a circular shape to an ovalshape) is prevented. Since the unison ring assembly 46 has goodout-of-plane stiffness, the number of projections 62 can be minimised.

FIG. 9 is a plan view of a pair of adjacent linkages 48, with the unisonring 52 located at a first position. In the configuration shown in FIG.9, the unison ring 52 is positioned at a relatively downwards positionas viewed from FIG. 9 and a relatively anti-clockwise position as viewedfrom the right-hand side of FIG. 4. In the configuration shown in FIG.9, the distance between the coupling between the vane 50 and the firstlink 72 and the coupling between the second link 74 and the unison ring52 (i.e. the distance between axes 51, 92) is relatively small.Accordingly, in order to accommodate the lengths of the first link 72and the second link 74, it is necessary for the first link 72 to bepivoted to a relatively anti-clockwise position as viewed from FIG. 9,such that a relatively small angle θ₁ is formed between the first link72 and the second link 74 and a relatively large angle is formed betweenthe first link 72 and the longitudinal axis 41.

FIG. 10 is a plan view of a pair of adjacent linkages 48, with theunison ring 52 located at a second position. In the configuration shownin FIG. 10, the unison ring 52 is positioned at a relatively upwardsposition as viewed from FIG. 10 (as shown by the difference between thepositions of the connecting rods 64 in FIGS. 9 and 10) and a relativelyclockwise position as viewed from the right-hand side of FIG. 4. In theconfiguration shown in FIG. 10, the distance between the couplingbetween the vane 50 and the first link 72 and the coupling between thesecond link 74 and the unison ring 52 (i.e. the distance between axes51, 92) is relatively large. Accordingly, in order to accommodate thelengths of the first link 72 and the second link 74, it is necessary forthe first link 72 to be pivoted to a relatively clockwise position asviewed from FIG. 10 such that a relatively large angle θ₂ is formedbetween the first link 72 and the second link 74 and a relatively smallangle is formed between the first link 72 and the longitudinal axis.

As mentioned previously, the first link 72 is secured to the vane 50such that they are unable to move relative to each other. Accordingly,when the unison ring 52 is positioned as shown in FIG. 9, the vane 50 isrotated to a relatively anti-clockwise position as viewed from FIG. 9,thereby reducing the flow area through the compressor. Likewise, whenthe unison ring 52 is positioned as shown in FIG. 10, the vane 50 isrotated to a relatively clockwise position as shown in FIG. 10, therebyincreasing the flow area through the compressor. Accordingly, byactuating the actuator and thus the position of the unison ring assembly46, the angle of the vane 50 and the performance of the compressor canbe changed.

The desired range of motion of the vane 50 may change before or aftermanufacture of the gas turbine engine 10. In such circumstances, therange of motion of the vane 50 may be changed simply by changing thelength of the first and/or second link 72, 74. This has minimal effecton surrounding components within the casing assembly 42 and gas turbineengine 10 as a whole. Accordingly, the need for significant redesign ofthe casing assembly 42 or the gas turbine engine 10 as a whole isprevented. In some circumstances, it may be preferable to change thelength of the second link 74 whilst keeping the length of the first link72 the same. In other circumstances, it may be preferable to change thelength of the first link 72 whilst keeping the length of the second link74 the same. The range of motion of the vane 50 may alternatively oradditionally be modified by changing the distance between the unisonring 52 and the axis around which the vane 50 rotates (i.e. the axialdistance between the axes 51 and 92). The lengths of the first andsecond links 72, 74 and the distance between the unison ring 52 and theaxis around which the vanes 50 rotate may be different for different rowof vanes 50 so as to effect different amounts of rotation of the vanes50 within different rows.

By way of example, in a casing assembly 42 having a first sizing, thelength of the first link 72 is 50 millimetres, the length of the secondlink 74 is 35 millimetres and the distance between the axis along whichthe unison ring 52 rotates and the axis about which the vane 50 rotatesis 45 millimetres. This results in a total range of unison ring 52travel of approximately 37 millimetres and a range of motion of the vane50 of approximately 48 degrees. In a casing assembly 42 having a secondsizing, the length of the first link 72 is 50 millimetres, the length ofthe second link 74 is 70 millimetres and the distance between the axisalong which the unison ring 52 rotates and the axis about which the vane50 rotates is 80 millimetres. This results in a total range of unisonring 52 travel of approximately 37 millimetres and a range of motion ofthe vane 50 of approximately 37 degrees.

The linkages 48 act as straight line linkages or straight linemechanisms. The linkages 48 allow the unison ring assembly 46 to berotated in a circumferential direction (i.e. around the longitudinalaxis 41) without moving axially (i.e. along the longitudinal axis 41).Since each of the unison rings 52, 54, 56 do not move axially, they canbe connected to form a single unison ring assembly 46 that can becontrolled directly by a single actuator. Accordingly, there is no needfor the unison rings 52, 54, 56 to be controlled individually, such asvia a plurality of control rods connected to a crankshaft, as istypically done in existing casing assemblies. The need for a pluralityof control rods and crankshaft is therefore avoided. As mentioned above,the lengths of the first and second links 72, 74 and the distancebetween the unison ring 52 and the axis around which the vanes 50 rotatemay be different for different rows of vanes 50, such that rotation ofthe unison ring assembly 46 effects different amounts of rotation indifferent vanes 50.

By reducing the number of components relative to existing casingassemblies, the size of the casing assembly 42 can be reduced, therebyallowing the casing assembly 42 to fit within existing gas turbineengines 10, as well as gas turbine engines having less available spacefor the casing assembly 42. Further, by both reducing the number ofcomponents and providing multiple components that are identical orsubstantially identical to each other (e.g. the first, second and thirdunison rings 52, 54, 56 and the connecting rods 64), the casing assembly42 is simple, quick and inexpensive to manufacture.

An additional benefit of providing a casing assembly 42 that does notmove axially is that the connection between the unison ring assembly 46and the casing 44 can be relatively simple in comparison to existingdesigns. One example of such a simple connection has already beendescribed with reference to FIG. 8. Alternative simple connectionsbetween unison ring assemblies and casings may alternatively beprovided, such as that shown in FIGS. 11 and 12.

FIGS. 11 and 12 are cross-sectional views of a connection between analternative unison ring 104 and an alternative casing 106. Theorientation of the cross-sectional plane of FIG. 11 corresponds that ofFIG. 8. The cross-sectional plane of FIG. 12 is perpendicular to thecross-sectional plane of FIG. 11. The alternative unison ring 104substantially corresponds to the unison ring 52. However, a first pin108 and a second pin 110 extend from a distal end (i.e. a radially innerend) of the projection 62. The alternative casing 106 comprises analternative first flange 112 and an alternative second flange 114 thatare spaced apart from each other so as to define a groove 116. As bestshown in FIG. 12, the alternative first and second flanges 112, 114 onlyextend part way around the circumference of the alternative casing 106.The alternative first and second flanges 112, 114 are provided with afirst slot 118 and a second slot 120, respectively. The first and secondpins 108, 110 extend through and are supported by the first and secondslots 118, 120, respectively. In alternative arrangements one or morerollers could run along the first and second slots 118, 120, rather thanthe first and second pins 108, 110.

Although it has been described that the connection between the secondlink 74 and the first link 72 is a hinge joint and that the connectionbetween the second link 74 and the unison ring 52 is a universal joint,in alternative arrangements both the connection between the second link74 and the first link 72 and the connection between the second link 74and the unison ring 52 may be universal joints. This may increase theamount by which the unison ring assembly 46 is able to rotate about thelongitudinal axis 41 upon actuation by the actuator. In furtheralternative arrangements, the connection between the second link 74 andthe first link 72 may be a universal joint and the connection betweenthe second link 74 and the unison ring 52 may be a hinge joint. In yetfurther alternative arrangements, both of the connections may be hingejoints. In yet further alternative arrangements, one or both of theconnections may be spherical bearings (i.e. ball joints).

Although it has been described that the casing assembly 42 is a casingassembly of a high pressure compressor, it may alternatively be a casingassembly of a low pressure compressor such as the low pressurecompressor 14 described with reference to FIGS. 1 and 2.

Although it has been described that three unison rings are provided forthree rows of vanes, any number of unison rings may be provided for anynumber of rows of vanes. For example, four unison rings may be providedfor four rows of vanes. Such an arrangement may be used in a highpressure compressor 15 such as that described with reference to FIGS. 1and 2, which typically comprise four rows of vanes. By way of furtherexample, a unison ring assembly comprising a single unison ring may beprovided for a single row of vanes, or a unison ring assembly comprisingfive unison rings may be provided for five rows of vanes.

Although it has been described that the first link 72 and the vane 50are separate components, in alternative arrangements they may be formedintegrally with each other. In such arrangements, as with the previouslydescribed arrangements, the first link comprises an elongate bodyextending away from the main body of the vane 50 (i.e. the aerofoilportion 70 of the vane 50) and the rotational axis 51 thereof.

Although the structure and operation of a single vane 50 andcorresponding portions of the rest of the casing assembly 42 have beendescribed, the remaining vanes and corresponding portions of the rest ofthe casing assembly 52 may be structured and operate in a similarmanner.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

What is claimed is:
 1. A casing assembly for a gas turbine engine, thecasing assembly comprising: an annular casing having a longitudinalaxis; a unison ring assembly configured for rotation about thelongitudinal axis; a plurality of vanes rotatably coupled to the annularcasing; and a plurality of linkages, wherein each of the plurality ofvanes is coupled to the unison ring assembly by a respective one of theplurality of linkages, wherein each linkage comprises: a first linkcoupled to the respective vane; and a second link rotatably coupled tothe first link and rotatably coupled to the unison ring assembly, suchthat rotation of the unison ring assembly about the longitudinal axiseffects rotation of the respective vane.
 2. The casing assembly asclaimed in claim 1, wherein the unison ring assembly is configured formovement on a single plane.
 3. The casing assembly as claimed in claim1, wherein the second link of each of the linkages is rotatably coupledto its respective first link about a first axis of rotation that issubstantially perpendicular to the longitudinal axis.
 4. The casingassembly as claimed in claim 1, wherein the second link of each of thelinkages is rotatably coupled to the unison ring assembly about a secondaxis of rotation that is substantially perpendicular to the longitudinalaxis.
 5. The casing assembly as claimed in claim 1, wherein the firstlink of each of the linkages is fixedly coupled to its respective vane.6. The casing assembly as claimed in claim 1, wherein the second link ofeach of the linkages is rotatably coupled to its respective first linkand/or the unison ring assembly by a hinge joint.
 7. The casing assemblyas claimed in claim 1, wherein the second link of each of the linkagesis rotatably coupled to its respective first link and/or the unison ringassembly by a universal joint.
 8. The casing assembly as claimed inclaim 1, wherein the second link of each of the linkages is rotatablycoupled to its respective first link and/or the unison ring assembly bya spherical joint.
 9. The casing assembly as claimed in claim 1, whereintwo or more of the first links or two or more of the second links havetwo or more different lengths.
 10. The casing assembly as claimed inclaim 1, wherein the unison ring assembly comprises two or more unisonrings spaced from each other, wherein the second links of each of thelinkages are rotatably coupled to the two or more unison rings.
 11. Thecasing assembly as claimed in claim 10, wherein the two or more unisonrings are connected by one or more connecting rods.
 12. The casingassembly as claimed in claim 10, wherein each of the two or more unisonrings comprise a plurality of clevises, wherein the second links of eachof the linkages are rotatably coupled to the unison ring assembly via arespective one of the plurality of clevises.
 13. The casing assembly asclaimed in claim 10, wherein each of the two or more unison ringscomprise a plurality of projections extending towards the longitudinalaxis, wherein the annular casing comprises a plurality of grooves orslots for receiving the plurality of projections and locating the two ormore unison rings relative to the annular casing.
 14. The casingassembly as claimed in claim 10, wherein the two or more unison ringsare substantially planar.
 15. The casing assembly as claimed in claim 1,wherein the plurality of vanes are a plurality of stator vanes and/or aplurality of inlet guide vanes.
 16. A gas turbine engine comprising acasing assembly as claimed in claim 1.